Reducing noise from a combustor of a gas turbine engine

ABSTRACT

A method of reducing noise from a combustor of a gas turbine engine includes the steps of establishing a maximum noise limit that may be for a particular frequency range. A primary fuel flow percentage, which may be emitted from a fuel nozzle arrangement having various groupings of simplex and duplex nozzles, is then established. An immersion depth measured between an aft rim of a swirler and a distal tip of the fuel nozzles may then be reduced thereby reducing the noise amplitude.

FIELD

The present disclosure relates to a combustor of a gas turbine engineand, more particularly, to a method of reducing noise from thecombustor.

BACKGROUND

The present disclosure relates to a combustor of a gas turbine engineand, more particularly, to a method of reducing noise from thecombustor.

Gas turbine engines, such as those that power modern commercial andmilitary aircraft, include a fan section to propel the aircraft,compressor section to pressurize a supply of air from the fan section, acombustor section to burn a hydrocarbon fuel in the presence of thepressurized air, and a turbine section to extract energy from theresultant combustion gases and generate thrust.

It remains desirable for gas turbine engine manufacturers to developcombustor configurations that reduce emissions and/or noise withimproved operational efficiencies.

SUMMARY

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

A combustor is disclosed. The combustor may have a fuel nozzle. The fuelnozzle may have a distal tip. The distal tip may include a tubularportion, a distal portion, and a conical portion concentricallyorientated and axially disposed between the tubular portion and distalportion. The distal tip may include a plurality of fuel spray aperturescircumferentially spaced from one another and communicating through aconical portion and configured to produce a fuel spray into a combustionchamber. The distal tip may include an immersion depth measured betweenthe distal portion of the distal tip and the aft most rim of theswirler, wherein the immersion depth is less than 0.500 inches. Invarious embodiments, the fuel spray apertures are angled at leastpartially radially outward and at least partially axially along acenterline of the fuel nozzle. Moreover, the fuel nozzle may include adistal aperture communicating through a planar portion of the distal tipand configured to produce a fuel spray along the centerline of the fuelnozzle.

A method of designing a low noise turbine engine combustor is disclosed.The method includes establishing a maximum noise limit, establishing aprimary fuel flow percentage, and reducing an immersion depth measuredbetween an aft rim of a swirler and a distal tip of a fuel nozzle untilthe noise generated drops below the maximum noise limit. The methodfurther includes wherein the primary fuel flow percentage is establishedfor low engine power conditions. The primary fuel flow percentage may beabout less than twenty-five percent.

The method may also include a step of increasing the primary fuel flowpercentage to further reduce combustor noise. The combustor may have anarrangement of simplex and duplex nozzles. In various embodimentsreducing the immersion depth at least partially desensitizes thecombustor against other noise producing factors including primary fuelflow. In various embodiments the immersion depth is less than 0.500inches. In various embodiments, the immersion depth is less than 0.350inches. Moreover, in various embodiments, the noise limit may be equalto or less than one percent of a combustor amplitude. The primary fuelflow may come from the duplex nozzles.

A method of reducing noise from a combustor of a gas turbine engineoperating at low power conditions is disclosed. The method may includemaximizing primary fuel flow and minimizing an immersion depth measuredbetween an aft rim of a swirler and a distal tip of a fuel nozzle. Invarious embodiments, a primary fuel flow percentage is about less thantwenty-five percent. In various embodiments, the combustor has anarrangement of simplex and duplex nozzles. Reducing the immersion depthmay at least partially desensitize the combustor against other noiseproducing factors including primary fuel flow. In various embodimentsthe immersion depth is less than 0.500 inches. In various embodiments,the combustor noise is measured for tones falling between a range ofabout 200 Hz to 400 Hz. A maximum noise limit may be equal to or lessthan one percent of a combustor amplitude. The primary fuel flow mayflow from the duplex nozzles.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed, non-limiting,embodiments. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is a cross-section of a combustor section;

FIG. 3 is a partial cross section of duplex fuel nozzle tip of thecombustor section;

FIG. 4 is a partial cross section of a simplex fuel nozzle tip of thecombustor section;

FIG. 5 is a cross section of the combustor section illustrating anarrangement of fuel nozzles taken in the direction of arrows 5-5 in FIG.2 ;

FIG. 6 is a cross section of the combustor section illustrating a secondembodiment of an arrangement, and taken in the direction of arrows 6-6in FIG. 2 ; and

FIG. 7 is an enlarged, partial, cross section of the combustor section.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration. While these exemplary embodiments are described insufficient detail to enable those skilled in the art to practiceembodiments of the disclosure, it should be understood that otherembodiments may be realized and that logical changes and adaptations indesign and construction may be made in accordance with this inventionand the teachings herein. Thus, the detailed description herein ispresented for purposes of illustration only and not limitation. Thescope of the disclosure is defined by the appended claims. For example,the steps recited in any of the method or process descriptions may beexecuted in any order and are not necessarily limited to the orderpresented. Furthermore, any reference to singular includes pluralembodiments, and any reference to more than one component or step mayinclude a singular embodiment or step. Also, any reference to attached,fixed, connected or the like may include permanent, removable,temporary, partial, full and/or any other possible attachment option.Additionally, any reference to without contact (or similar phrases) mayalso include reduced contact or minimal contact.

Furthermore, any reference to singular includes plural embodiments, andany reference to more than one component or step may include a singularembodiment or step. Surface shading lines may be used throughout thefigures to denote different parts but not necessarily to denote the sameor different materials.

As used herein, “aft” refers to the direction associated with theexhaust (e.g., the back end) of a gas turbine engine. As used herein,“forward” refers to the direction associated with the intake (e.g., thefront end) of a gas turbine engine.

In accordance with various aspects of the disclosure, apparatuses,systems and methods are described for providing a distributed fuelinjection in connection with an aircraft engine. In various embodiments,a gas turbine engine may exhibit a tendency to generate tones (e.g.,sound having one or more frequency and/or amplitude). For instance, thefuel exiting fuel nozzles may form a film on a swirler surface with athickness or other properties that allows airflow perturbations tocouple with fuel film causing unsteady heat release which can thenamplify the natural frequency of the combustor causing a tone. Also, thefuel exiting fuel nozzles may insufficiently couple with an airflow sothat the lack of mixing and travel to the swirler surface inhibitsanchoring of a flame to the surface, causing flame stability issueswhich can couple with a natural frequency of the combustor and alsocause tones.

The combustor section may have an annular wall assembly having inner andouter shells that support respective inner and outer heat shieldingliners. The liners may be comprised of a plurality of floating heatshields or panels that together define an annular combustion chamber. Anannular cooling cavity is defined between the respective shells andliners for supplying cooling air to an opposite hot side of the panelsthrough a plurality of strategically placed effusion holes. Impingementholes are located in the shell for supply cooling air from an outer airplenum and into the cavity. The effusion holes are generally orientatedto create a protective blanket, or, air film over the hot side of thepanels, thereby protecting the panels from the hot combustion gases inthe chamber.

A forward assembly or bulkhead of the combustor generally supports aplurality of circumferentially spaced swirlers that each receive a fuelnozzle that projects rearward and toward the combustion chamber. Fuel iscontrollably supplied by each nozzle and compressed air generally flowsbetween the swirler and nozzle and into the combustion chamber. Thecombination of the nozzle and swirler facilitates mixing of the fuel andair for combustion in the chamber.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbo fan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith turbofans as the teachings may be applied to other types of turbineengines such as a turbojets, turboshafts, and three spool (plus fan)turbofans wherein an intermediate spool includes an intermediatepressure compressor (“IPC”) between a low pressure compressor (“LPC”)and a high pressure compressor (“HPC”), and an intermediate pressureturbine (“IPT”) between the high pressure turbine (“HPT”) and the lowpressure turbine (“LPT”).

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about a central, longitudinal, engine axis Arelative to an engine static structure 36 or engine case via severalbearing structures 38. The low spool 30 generally includes an innershaft 40 that interconnects a fan 42 of the fan section 22, a LPC 44 ofthe compressor section 24 and a LPT 46 of the turbine section 28. Theinner shaft 40 drives the fan 42 directly or through a gearedarchitecture 48 to drive the fan 42 at a lower speed than the low spool30. An exemplary reduction transmission is an epicyclic transmission,namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a HPC 52of the compressor section 24 and HPT 54 of the turbine section 28. Acombustor 56 of the combustor section 26 is arranged between the HPC 52and the HPT 54. The inner shaft 40 and the outer shaft 50 are concentricand rotate about the engine central longitudinal axis A that iscollinear with their longitudinal axes. Core airflow is compressed bythe LPC 44 then the HPC 52, mixed with the fuel and burned in thecombustor 56, then expanded over the HPT 54 and the LPT 46. The LPT 46and HPT 54 rotationally drive the respective low spool 30 and high spool32 in response to the expansion.

In various non-limiting examples, the gas turbine engine 20 is ahigh-bypass geared aircraft engine. In a further example, the gasturbine engine 20 bypass ratio is greater than about six (6:1). Thegeared architecture 48 can include an epicyclic gear train, such as aplanetary gear system or other gear system. The example epicyclic geartrain has a gear reduction ratio of greater than about 2.3:1, and inanother example is greater than about 2.5:1. The geared turbofan enablesoperation of the low spool 30 at higher speeds that can increase theoperational efficiency of the LPC 44 and LPT 46 and render increasedpressure in a fewer number of stages.

A pressure ratio associated with the LPT 46 is pressure measured priorto the inlet of the LPT 46 as related to the pressure at the outlet ofthe LPT 46 prior to an exhaust nozzle of the gas turbine engine 20. Inone non-limiting embodiment, the bypass ratio of the gas turbine engine20 is greater than about ten (10:1), the fan diameter is significantlylarger than that of the LPC 44, and the LPT 46 has a pressure ratio thatis greater than about five (5:1). It should be understood; however, thatthe above parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans.

In one embodiment, a significant amount of thrust is provided by thebypass flow path B due to the high bypass ratio. The fan section 22 ofthe gas turbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet(10,668 meters). This flight condition, with the gas turbine engine 20at its best fuel consumption, is also known as thrust specific fuelconsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan pressure ratio is the pressure ratio across a blade of the fansection 22 without the use of a fan exit guide vane system. The low fanpressure ratio according to one, non-limiting, embodiment of the examplegas turbine engine 20 is less than 1.45. low corrected fan tip speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (T/518.7)0.5 in which “T” represents the ambienttemperature in degrees Rankine. The low corrected fan tip speedaccording to a non-limiting embodiment of the example gas turbine engine20 is less than about 1,150 feet per second (351 meters per second).

Referring to FIG. 2 , the combustor section 26 generally includes acombustor 56 with an outer combustor wall assembly 60, an innercombustor wall assembly 62 and a diffuser case module 64 that encasesassemblies 60, 62. The outer combustor wall assembly 60 and the innercombustor wall assembly 62 are radially spaced apart such that anannular combustion chamber 66 is defined therebetween.

The outer combustor wall assembly 60 is spaced radially inward from anouter diffuser case 64A of the diffuser case module 64 to define anouter annular plenum 76. The inner combustor wall assembly 62 is spacedradially outward from an inner diffuser case 64B of the diffuser casemodule 64 to define an inner annular plenum 78. It should be understoodthat although a particular combustor is illustrated, other combustortypes with various combustor liner arrangements will also benefit. Itshould be further understood that the disclosed cooling flow paths arebut an illustrated embodiment and should not be limited.

The combustion chamber 66 contains the combustion products that flowaxially toward the turbine section 28. Each combustor wall assembly 60,62 generally includes a respective support shell 68, 70 that supportsone or more liners 72, 74 mounted thereto. Each of the liners 72, 74 maybe formed of a plurality of floating heat shields or panels that aregenerally rectilinear and manufactured of, for example, a nickel basedsuper alloy that may be coated with a ceramic, or other temperatureresistant material, and are arranged to form a liner array. Each linermay have a plurality of forward panels 72A and a plurality of aft panels72B that line the outer shell 68. A plurality of forward panels 74A anda plurality of aft panels 74B also line the inner shell 70. At least oneigniter 79 is generally located at and projects through the forwardpanel 72A to initially ignite a blended fuel-air mixture. It should beappreciated that the liner array may alternatively include but a singlepanel rather than the illustrated axial forward and axial aft panels.

The combustor 56 further includes a forward assembly 80 immediatelydownstream of the compressor section 24 to receive compressed airflowtherefrom. The forward assembly 80 generally includes an annular hood82, a bulkhead assembly 84, and a plurality of swirlers 90 (one shown)that are circumferentially spaced from one-another. Each swirler 90 iscircumferentially aligned with one of a plurality of fuel nozzles 86(one shown) and a respective one of a plurality of hood ports 94. Thebulkhead assembly 84 includes a bulkhead support shell 96 secured to thecombustor walls 60, 62, and a liner that may have a plurality ofcircumferentially distributed bulkhead heat shields or panels 98 securedto the bulkhead support shell 96 around each of a respective swirleropening 92. The bulkhead support shell 96 is generally annular and theplurality of circumferentially distributed bulkhead panels 98 aresegmented, typically one to each fuel nozzle 86 and swirler 90.

The annular hood 82 extends radially between, and is secured to, theforwardmost ends of the combustor wall assemblies 60, 62. Each one ofthe plurality of circumferentially distributed hood ports 94 receives arespective one of the plurality of fuel nozzles 86 and facilitates thedirection of compressed air into the forward end of the combustionchamber 66 through the swirler opening 92. Each fuel nozzle 86 may besecured to the diffuser case module 64 and projects through one of thehood ports 94 into the respective swirler opening 92.

The forward assembly 80 introduces core combustion air into the forwardsection of the combustion chamber 66 while the remainder enters theouter annular plenum 76 and the inner annular plenum 78. The pluralityof fuel nozzles 86 and adjacent structure generate the blended fuel-airmixture that supports stable combustion in the combustion chamber 66.

Referring to FIGS. 3 and 4 , the forward assembly 80 may utilize avariety of structurally different fuel nozzles 86 for varying thefuel-air mixture at different circumferential locations within theannular combustion chamber 66 to control fuel burn, temperature profilesand combustor noise. For example, the plurality of fuel nozzles 86 mayinclude at least one duplex fuel nozzle 86A and at least one simplexfuel nozzle 86B. A distal tip 100 of the nozzle 86 generally projectsalong a centerline C, and includes a generally tubular portion 102, adistal or planar portion 104, and a conical or frustum portion 106concentrically orientated and axially disposed between the tubular anddistal portions 102, 104.

Both of the duplex and simplex fuel nozzles 86A, 86B may have aplurality of fuel spray apertures or orifices 108 communicating throughthe conical portion 106 of the distal tip 100 for producing a secondaryfuel spray (see secondary fuel spray 110) generally into the combustionchamber 66. The apertures 108 may be circumferentially spaced from oneanother and are generally angled at least partially radially outwardfrom, and also at least partially axially along the centerline C. Invarious embodiments, the aperture 108 may be singular and generallyannular or circular in shape (i.e. circumferentially continuous). Theduplex fuel nozzle 86A may further have at least one distal aperture ororifice 112 communicating through the planar portion 104 of the tip 100for discharging a cone-like, primary, fuel spray (see arrow 114) in anaxial direction with respect to centerline C. In contrast, the simplexfuel nozzle 86B may have no apertures in the distal portion 104.

Referring to FIGS. 3 through 5 , in various embodiments, fuel nozzlearrangement 116A is illustrated having eighteen fuel nozzles 86 witheach nozzle being circumferentially spaced from the next adjacentnozzle, and the spacing between nozzles may be substantially equal tothe next spacing. Arrangement 116A may have four nozzle groupings 118A,120A, 122A, 124A aligned circumferentially with one another. Grouping118A may be circumferentially located between groupings 120A, 124A, andmay have three simplex nozzles 86B. Grouping 120A may becircumferentially located between groupings 118A, 122A, and may have sixsimplex nozzles 86B. Grouping 122A may be circumferentially locatedbetween groupings 120A, 124A, and may have three duplex nozzles 86A.And, grouping 124A may be circumferentially located between groupings118A, 122A, and may have six simplex nozzles 86B.

The primary fuel flow or spray 114 may be supplied to the duplex nozzles86A via a primary manifold 126 and the secondary fuel flow or spray 110may be supplied to the duplex and simplex nozzles 86A, 86B via asecondary manifold 128. Fuel supplied to the manifolds 126, 128 and/orfuel pressure within the manifolds may be controlled through any varietyof means typically known in the art. During low power operatingconditions of the engine 20, the primary manifold 126 may be providedwith a higher fuel pressure than the secondary manifold 128 to drive afuel flow distortion. The increased fuel pressure drop (and fuel flow)increases the overall (primary plus secondary) fuel flow (for instance,mass over time and/or volume over time) to the duplex nozzles 86Arelative to the simplex nozzles 86B. Thus, the duplex nozzles 86Agenerate the relatively high fuel-air ratio mixtures (relative to thesimplex nozzles 86B) and the simplex nozzles 86B provide the relativelylow fuel-air ratio mixtures (relative to the duplex nozzles 86A). Thevaried fuel-air ratio mixtures may partially dampen tangential and axialpressure waves within the combustor 56 to, in-part, control combustortones and enhance combustor stability. At low power conditions, fuelsupply pressure may generally be around 100 psi (689.5 kPa), with anapproximate 50 to 100 psi (about 344.7 to about 1034 kPa) differencebetween the duplex and simplex nozzles 86A, 86B. It should beappreciated that “low power” as defined herein may include moderatepower such as that required for approach conditions and a margin aboveat least partially into a cruise condition.

At non-low power operating conditions of the engine 20, fuel pressuremay be about 1200 psi (about 8273 kPa) and the primary and secondarymanifolds 126, 128 may receive equalized flow such that the duplex andsimplex nozzles 86A, 86B generate a symmetrically uniform fuel-air ratiothroughout the combustor 56. It should be appreciated that “non-lowpower” as defined herein may include cruise power conditions and abovesuch as a take-off flight condition.

Referring to FIGS. 3, 4 and 6 , another, non-limiting, example of a fuelnozzle arrangement 116B is illustrated having eighteen fuel nozzles 86with each nozzle being circumferentially spaced from the next adjacentnozzle, and the spacing between nozzles may be substantially equal tothe next spacing. Arrangement 116B may have four nozzle groupings 118B,120B, 122B, 124B aligned circumferentially with one another. Grouping118B may be circumferentially located between groupings 120B, 124B, andmay have six duplex nozzles 86A. Grouping 120B may be circumferentiallylocated between groupings 118B, 122B, and may have three simplex nozzles86B. Grouping 122B may be circumferentially located between groupings120B, 124B, and may have six duplex nozzles 86A. And, grouping 124B maybe circumferentially located between groupings 118B, 122B, and may havethree simplex nozzles 86B. It is further contemplated and understoodthat either or both arrangements 116A, 116B may have a different numberof fuel nozzles 86 than illustrated, with a different number ofgroupings that may be characterized by different types of fuel nozzlesand not necessarily simplex and duplex designs.

Referring to FIGS. 3 and 7 , the distal tip 100 of each fuel nozzle 86(a duplex fuel nozzle 86A being illustrated) projects along thecenterline C and into a surrounding, cone-shaped, cavity 130 defined bythe swirler 90 that converges toward the centerline C is a downstreamdirection. The cavity 130 is in direct fluid communication with andlocated upstream of the combustion chamber 66. The distal tip 100deposits the secondary and primary fuel sprays 110, 114 into the cavity130 where the fuel initially mixes with the compressed air and isexpelled into the combustion chamber 66 for continued mixing andcombustion.

The swirler 90 spans axially with respect to centerline C and betweenforward and aft (i.e. upstream and downstream) rims 132, 134 of theswirler. The nozzle 86 is axially orientated to the swirler 90 by animmersion depth (see immersion depth 136) measured between the distalportion 104 of the nozzle tip 100 and the downstream (e.g., the aftmost, relative to the centerline C) rim 134 of the swirler 90.

In accordance with the principles discussed herein, combustor noise, atlow power operating conditions, can be minimized by reducing immersiondepth 136 and/or by increasing the percent of primary fuel flow 114. Invarious embodiments, the significant difference between the arrangements116A, 116B (as two, non-limiting examples), is an indication that most,if not all, other arrangements of various grouping of duplex and simplexnozzles 86A, 86B will generally follow the same noise abatement trends.For instance, an immersion depth 136 (i.e. measured between the swirler90 and the nozzle tip 100), versus combustor noise amplitude may followa trend. In various embodiments, the immersion depth 136 may be selectedfrom between about 0.250 inches to about 0.500 inches (6.35 mm to 12.70mm), where the term “about” in this context only may refer to +/−0.05inches. Moreover, the immersion depth may be expressed as a ratio ofimmersion depth over swirler axial length. A maximum goal limit forcombustor noise amplitude may be generally dictated by governmentalregulations and the immersion depth may be selected to meet a specifiedregulation. In various embodiments, an immersion depth comprises lessthan or equal to about 0.500 inches (12.70 mm), where the term “about”in this context only may refer to +/−0.05 inches.

Moreover, in various embodiments, such as for an arrangement 116A havinga primary fuel flow 114 of about 2.5% of total flow, an immersion depth136 of about 0.350 inches (8.89 mm) or less may be desirable. Moreover,for arrangement 116A having a primary fuel flow 114 of about 5% of totalflow, an immersion depth 136 of about 0.360 inches (9.14 mm) or less maybe desirable. Moreover, for the arrangement 116B having a primary fuelflow 114 of about 15%, an immersion depth 136 of about 0.390 inches(9.91 mm) or less may be desirable. Similarly, for an arrangement 116Bhaving a primary fuel flow 114 of about 25%, it may be generallydesirable to maintain an immersion depth 136 of about 0.500 inches(12.70 mm) or less.

Furthermore, the engine speed verse combustor amplitudes for tonesranging from about 200 Hz to 400 Hz may be of relevance to selectingimmersion depths. In general, engine low power may fall within a rangeof about 2000 to 3000 rpm; engine mid power may fall within a range ofabout 3000 to 6500 rpm; and engine high power is generally denoted by6500 rpm's or higher. At engine low power conditions for tones of 200 Hzto 400 Hz, combustor amplitude may be reduced by a factor such as aboutfifteen times by decreasing immersion depth and increasing percent ofprimary fuel flow according to the principles herein. At engine midpower conditions, combustor amplitude may be increased by a factor ofabout two, and at engine high power conditions combustor amplitude issubstantially the same.

Still furthermore, the engine speed verse combustor amplitudes for tonesranging from about 400 Hz to 800 Hz may also be of relevance toselecting immersion depths. At engine low power and mid power conditionsfor tones of 400 Hz to 800 Hz, combustor amplitudes may remainsubstantially such that decreasing immersion depth and increasing thepercent of primary fuel flow has little effect. However, at engine highpower conditions combustor amplitude may decrease by a factor of abouttwelve when the percent of primary fuel flow 114 is increased and theimmersion depth 136 is decreased.

While the disclosure is described with reference to exemplaryembodiments, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted withoutdeparting from the spirit and scope of the disclosure. In addition,different modifications may be made to adapt the teachings of thedisclosure to particular situations or materials, without departing fromthe essential scope thereof. The disclosure is thus not limited to theparticular examples disclosed herein, but includes all embodimentsfalling within the scope of the appended claims.

What is claimed is:
 1. A combustor comprising a plurality of swirlersthat are circumferentially spaced from one another and a correspondingplurality of fuel nozzles such that each swirler of the plurality ofswirlers is axially aligned with a different fuel nozzle of theplurality of fuel nozzles, the plurality of fuel nozzles comprising aplurality of duplex fuel nozzles and a plurality of simplex fuelnozzles, the duplex fuel nozzles being structurally different from thesimplex fuel nozzles, each fuel nozzle of the plurality of fuel nozzlescomprising: a distal tip, wherein the distal tip comprises: a tubularportion; a distal portion; and a conical portion concentricallyorientated and axially disposed between the tubular portion and thedistal portion; a plurality of fuel spray apertures circumferentiallyspaced from one another and communicating through the conical portionand configured to produce a fuel spray into a combustion chamber, thedistal portion projecting in a direction of and being recessed from adownstream rim of the plurality of swirlers corresponding with theplurality of fuel nozzles, an immersion depth measured between thedistal portion of the distal tip and the downstream rim, wherein theimmersion depth comprises less than 0.500 inches, wherein the combustorcomprises four nozzle groupings that are circumferentially disposedrelative to an axis of the combustor, wherein a first nozzle grouping ofthe nozzle groupings and a third nozzle grouping of the nozzle groupingseach comprise six of the plurality of duplex fuel nozzles, wherein asecond nozzle grouping of the nozzle groupings and a fourth nozzlegrouping of the nozzle groupings each comprise three of the plurality ofsimplex fuel nozzles, and that are circumferentially located between thefirst nozzle grouping and the third nozzle grouping.
 2. The combustor ofclaim 1, wherein the fuel spray apertures are angled at least partiallyradially outward and at least partially axially along a centerline ofthe plurality of fuel nozzles.
 3. The combustor of claim 1, wherein eachof the plurality of fuel nozzles comprises a distal aperturecommunicating through a planar portion of the distal tip and configuredto produce a fuel spray along a centerline of the plurality of fuelnozzles.
 4. The combustor of claim 1, wherein the first nozzle groupingis diametrically opposed to the third nozzle grouping, and wherein thesecond nozzle grouping is diametrically opposed to the fourth nozzlegrouping.
 5. The combustor of claim 1, wherein the second nozzlegrouping is adjacent to each of the first nozzle grouping and the thirdnozzle grouping on one side of the combustor, and wherein the fourthnozzle grouping is adjacent to each of the first nozzle grouping and thethird nozzle grouping on an opposite side of the combustor.
 6. Thecombustor of claim 1, wherein each swirler of the plurality of swirlersconverges relative to a centerline of the corresponding fuel nozzle ofthe plurality of fuel nozzles proceeding to the downstream rim, andwherein the downstream rim of each swirler of the plurality of swirlersdefines an open discharge end.
 7. A combustor comprising a plurality offuel nozzles comprising a distal tip, wherein the distal tip comprises:a tubular portion; a distal portion; and a conical portionconcentrically orientated and axially disposed between the tubularportion and the distal portion; a plurality of fuel spray aperturesspaced from one another and communicating through the conical portionand configured to produce a fuel spray into a combustion chamber of thecombustor, an immersion depth measured between the distal portion of thedistal tip and a downstream rim of a corresponding swirler of thecombustor, the distal portion projecting in a direction of and beingrecessed from a downstream rim of the corresponding swirler, wherein theimmersion depth is no more than about 0.500 inches, wherein thecombustor comprises a first nozzle grouping, a second nozzle grouping, athird nozzle grouping, and a fourth nozzle grouping, wherein the firstnozzle grouping and the third nozzle grouping each include a commonfirst number of a plurality of duplex nozzles of the plurality ofplurality of fuel nozzles, wherein the second nozzle grouping and thefourth nozzle grouping each include a common second number of aplurality of simplex nozzles of the plurality of fuel nozzles, whereinthe plurality of simplex nozzles are structurally different from theplurality of duplex nozzles, and wherein the first number is of adifferent magnitude than the second number.
 8. The combustor of claim 7,wherein the immersion depth is from between about 0.250 inches to about0.500 inches.
 9. The combustor of claim 7, wherein the first nozzlegrouping, the second nozzle grouping, the third nozzle grouping, and thefourth nozzle grouping are collectively disposed in a circumferentialarrangement.
 10. The combustor of claim 9, wherein the second nozzlegrouping is disposed between the first nozzle grouping and the thirdnozzle grouping on one side of the circumferential arrangement, andwherein the fourth nozzle grouping is disposed between the first nozzlegrouping and the third nozzle grouping on a different side of thecircumferential arrangement.
 11. The combustor of claim 10, wherein thefirst number for each of the first nozzle grouping and the third nozzlegrouping of the plurality of duplex nozzles is larger than the secondnumber for each of the second nozzle grouping and the fourth nozzlegrouping of the plurality of simplex nozzles.
 12. The combustor of claim10, wherein the first nozzle grouping and the third nozzle grouping aredisposed in opposing relation in the circumferential arrangement, andwherein the second nozzle grouping and the fourth nozzle grouping aredisposed in opposing relation in the circumferential arrangement. 13.The combustor of claim 12, wherein the first number for each of thefirst nozzle grouping and the third nozzle grouping of the plurality ofduplex nozzles is larger than the second number for each of the secondnozzle grouping and the fourth nozzle grouping of the plurality ofsimplex nozzles.
 14. The combustor of claim 9, wherein the first nozzlegrouping and the third nozzle grouping each include six duplex nozzlesof the plurality of duplex nozzles, and wherein the second nozzlegrouping and the fourth nozzle grouping each include three simplexnozzles of the plurality of simplex nozzles.
 15. The combustor of claim7, wherein the first number for each of the first nozzle grouping andthe third nozzle grouping of the plurality of duplex nozzles is largerthan the second number for each of the second nozzle grouping and thefourth nozzle grouping of the plurality of simplex nozzles.
 16. Thecombustor of claim 15, wherein the first nozzle grouping and the thirdnozzle grouping of the plurality of duplex nozzles are disposed inopposing relation to one another, and wherein the second nozzle groupingand the fourth nozzle grouping of the plurality of simplex nozzles aredisposed in opposing relation to one another.
 17. The combustor of claim7, wherein the first nozzle grouping and the third nozzle groupings eachinclude six duplex nozzles of the plurality of duplex nozzles, andwherein the second nozzle grouping and the fourth nozzle grouping eachinclude three simplex nozzles of the plurality of simplex nozzles. 18.The combustor of claim 17, wherein the first nozzle grouping and thethird nozzle grouping of the plurality of duplex nozzles are disposed inopposing relation to one another, and wherein the second nozzle groupingand the fourth nozzle grouping of the plurality of simplex nozzles aredisposed in opposing relation to one another.
 19. The combustor of claim7, wherein the corresponding swirler of the combustor converges relativeto a centerline of the fuel nozzle of the plurality of fuel nozzlesproceeding to the downstream rim, and wherein the downstream rim of thecorresponding swirler of the combustor defines an open discharge end.